A mathematical model for Vertical Attitude Takeoff and Landing (VATOL) aircraft simulation. Volume 3: User's manual for VATOL simulation program

APPENDIX h

DEFINITION AND IDCATION OF C - N PARAMETERS

W R TIIE V A W PROGRM

A. 1 fntaoducfidti This Appendix provides a l i s t i n g by coded n a e and defintlon of a l l V A W C m parameters arranged in order by segment and elemnt designation.

Each parameter is indicated 3 . type C o r N where C means t h e parameter is calculated by the program and doeo not need t o be quantified by program input and N mans t h e parameter is not calculated and m u s t be quantified by program input. Many of t h e p a r e t e r s in COMMON &re a l s o used in Volume I1 of t h i s report which describes the model equations and base a i r c r a f t data f o r VATOL

The Volume IK symbology for these parameters is included

simulation.

Note t h a t every parameter i n in parentheses following t h e definitions.

C O W N is a REAL program variable; therefore a l l non zero input data t o C O W N must have a decimal point.

A.2 Segment 1 - T i m e , P r i n t and P l o t Control Parameters

Coded Segment Element Type Name Units Definition sec Time which is updated every DT 1 1 C T seconds ( t ) 1 2 C TO sec T i m e controlled by RUNGE which takes on three intermediate values between T and WDT. Since

4th order Runge - Kutta integra-

t i o n is used, the basic integra- tion i n t e r v a l must be divided in four parts.

N DT sec Basic integraLion i n t e r v a l f o r RUNGE N RiAx sec Maximum value of T for a time h i s t o r y N DTPRNT sec I n t e r v a l between times of printed output I n t e r v a i between times of plotted

N D T P ~ T aec

output N TI sec I n i t i a l time for t i m e history; normally taken to be zero A-2

A . 3 Segment 2 - Variables t o be Integrated

Coded

wit8 Definition

Segment Element Type Nam3

2 1 C d T ftleec X body axis component of i n e r t i a l

acceleration ( i r )

f tlsec 2

2 2 C Y body axis component of i n e r t i a l

VOOT

acceleration ( 6 )

3 C Z body axis component of i n e r t i a l

2 f tlsec

WPQr

acceleration (i)

2 4 C Body axis r o l l acceleration (b)

P I ) @ r 8 4 h c 2

Body axis p i t c h a c c e l e r a t i o n ( 4 )

2 5 C radlsec

Q *

6 C rad/ sec Body axis yaw acceleration (i)

angT

2 7 C D l l D llsec Rate of change of dll d i r e c t i o n

cosine (a,,) 8 C Rate of Fhange of d12 d i r e c t i o n

2 D U D 11 sec

cosine (d12)

2 3 C D13D llsec Rate of change of d13 d i r e c t i o n

cosine (al3)

2 C D21D llsec Rate of change of dZ1 d i r e c t i o n

cosine (d ) 2 1

2 11 C D22D l i s e c Rate of change of dZ2 d i r e c t i o n

cosine (i22)

12 l l s e c Rate of change of dZ3 d i r e c t i o n 2 C D 2 3 D cosine 2 13 C D31D l l s e c Rate of Shange of d31 d i r e c t i o n cosine (d3r) F a t e of Shange of d d i r e c t i o n 2 1 4 C D32D l l s e c cosine (d32)

15 C D33D l l s e c Rate of change of d3,, d i r e c t i o n

cosine (a33)

2 16 C f t l e e c X i n e r t i a l axis component of

XEDdT

i n e r t i a l velocity (ie)

f t l s e c Y i n e r t i a l axis component of

2 17 C

YEWT i n e r t i a l velocity A-3 Coded

Segment Element Type Name Unite Dcf i a i t i o n

Z i n e r t i a l aria component of

18 C ZEWT f t / s e c

inertial velocity ( i , )

C LEUVND rad/eec f t elevon deflection rate

2 19

8 (1))

mFW

C RELKVND rad/sec Right elevon def l e c t l o n rate

2 20

(kp (2))

W

Rudder d e f l e c t i o n rate (gr)

2 21 C DRD rad/sec

C Normalized rate of change of-

2 22-23 WD(I) l / S e C afterburner thrust; 1 1 1 f o r l z f t engine, -2 f o r r i g h t engine ( T A B ( I ) ) PSITD(1) rad/sec Yaw t h r u s t deflection rate; 1-1

2 24-25 C

f q r l e f t engine, =2 f o r r i g h t engine

WT(l) 1

26-27 C THTD(1) rad/sec P i t c h t h r u s t deflection rate; 1-1

f q r l e f t engine, =2 f o r r i g h t engine

(e,(I) 1

C Normalized area deflection rate of 2 28-37 DRCSD(1) 1/ sec Xth RCS jet; space f o r 10 RCS , jets, only 2 used f o r VATLAS (&Rcs(I)) Rate of change of force of I t h R C S 2 38-47 C FRCSAD(1) lb/sec jet; space f o r 10-RCS jets, only 2 used f o r VATLAS (iRcsA(I)) Canard.trailing edge f l a p deflection 2 48 C DFLAPHD rad/sec

rate (4 )

TEFH

edge f l a p deflection 2 49 C D U W D rad/sec Normalized rate of change of non-

50-51 c TPPACD(1) llsec

2 .

afterburning t h r u s t ; 111 f o r ;eft

engine, -2 f o r r i g h t englne (TF(I))

52 C PCTRMD rad/eec Rate of change o f r o l l command

p5th integrator output ( f lgure 2 . 2 )

(PCW 1

PINTD rad/sec Body axis roll rate; a l s o coded a8

2 53 C P i n program @) A-4 Coded

S e w n t Eleman t Type Name Unite Dcf i n i t i o n

2 54 radleec Rate of change of r o l l forward

C PETRI0 path integrator output (figure 2.2)

(fie 1

TBM

rad/ sec Rate of change of yaw command path

55 C R c T m D

2 .

integrator output (figure 2.4)

(3 1

2 56 radleec Body axis yaw rate; a l s o coded as

C m T D

R in program (r)

2 rad/ se c Rate of change of yaw forward path 57 C RETRUD %tegrator output (figure 2 . 4 ) (reTm)

2 58 C YTRMD radlsec Rate of change of yaw trim input

(figure 2.4) (ETRM)

2 59 radlsec Rate of change of p i t c h conanand path

, c

Q C m

inthgrator output (figure 2.3)

2 60 rad/ sec Body axis p i t c h rate; also coded C

Qm

as Q in program ((I)

rad/sec 2 6 1 C Rate of change of p i t c h forward Q E m path i n t e g r a t o r ouaput: (figure 2.3) ( ; )

%RU

2 62 PTRMD radisec Rate of change of p i t c h t r i m input C

(figure 2.3) (4=)

2 63 C ZEDTRMD l b l s e c Rate of change of heave forward p t t h i n t e g r a t o r output (figure 2.5)

' T e * 1

2 64 C THETD rad 1 sec Rate of change of p i t c h angle ( 6 )

PHID radlsec Rate of change of r o l l angle (i)

2 65 C * 2 66 C PSID radlsec Rate of change of yaw angle (y?

2 C LNGSTKD

67 llsec Rate of change of p i t c h input shaping

f i l t e r output (i;)

2 68 C ROLLCD

llsec Rate of change ef r o l l input shaping

f i l t e r output (p;)

69 C Y A W 0 . l l e e c Rate of change of yaw input shaping

f i l t e r output (k;)

A-5

lfrPaant 3 - fatagratad Varirbtea

A.4 Coded

Segment Element Type Name Unite Definition

1 C U f t /sec X body axle component of inertial

v e l o c i t y (u)

2 V ft/.ec Y body axis component of inertial

3 C t e l o c i t y (v)

Z body axis component of i n e r t i a l

3 3 C w f t l s e c

velocity (w)

4 C P rad/ sec Body axis r o l l rate @)

3 5 C radlaec Body axis p i t c h rate (q)

Q

6 R radlsec Body axis yaw rate (r)

3 C L 3 7 C D 1 1 dll d i r e c t i o n cosine (dll) I d12 d i r e c t i o n cosine (d12) 3 8 C D12

-

C D 1 3 d13 d i r e c t i o n cosine (d13)

3 9

u 3 10 C D21 d21 d i r e c t i o n cosine (d21)

--

dz2 d i r e c t i o n cosine (dz2) 3 1 1 C D22

--

12 C D 2 3 d23 d i r e c t i o n cosine (d23) I 3 1 3 C D31 d31 d i r e c t i o n cosine (d31)

-

3 1 4 C D32 d32 direction cosine (d32)

--

C D3 3 d i r e c t i o n cosine (dJ3)

3 15

d33

16 C XE f t X position of cg i n i n e r t i a l

coordinates (xe)

C YE f t Y position of cg i n inertial

3 17 coordinates (ye)

3 18 C ZE f t Z position of cg in i n e r t i a l

coordinates (2,)

3 19 C LeLEvN rad ( I f t elevon d e f l e c t i o n (d (1))

TEFW

20 C RELEVN rad Right elevon deflection (d (2))

TEFW rad

3 21 C DR Rudder deflection ( Sr)

A-6 Coded

Segment Element Type Name wits De f i n 1 t ion

3 22-23 c T-U(I) - Normalized afterburner t h r u s t ; I =

1 f o r l e f t engine, -2 f o r r i g h t

engine ( T A B ( U 1

: 3 24-25 C PSIT(1) rad Yaw thrust deflection; 1-1 f o r

l e f t engine, -2 f o r r i g h t engine

(VT (1 11

3 26-27 C m(1) rad P i t c h t h r u s t deflection; 1-1 f o r l e f t engine, -2 f o r r i g h t engine 3 28-37 C DRCS(1) - Normalized area of I t h RCS jet; space f o r 10 RCS jets, only 2 used f o r VATLAS ( dRcs(I))

3 38-47 C FRCSA(1) l b Force of I t h RCS jet; space f o r

10 RCS jet8, only 2 used f o r VATLAS (FRa ) A

.3 48 C DFIAPH rad Canard t r a i l i n g edge f l a p d e f l e c t i o n

DFLAPU rad Wing leading edge f l a p deflection

3 49 C

%FW' 3 50-51 C TFRAC(1) -- Normalized non-af terburning t h r u s t ; 1=I f o r l e f t engine,=2 f o r r i g h t engine (TFI

3 52 C PCTRM rad R o l l command path i n t e g r a t o r output

(figure 2 - 2 1 (p, 1

TRM

3 53 C PINT rad I n t e g r a l of body axis r o l l r a t e

( f i g u r e 2.2) (pmT) 5 4 C PETRM rad R o l l forward path' i n t e g r a t o r output ( f i g u r e 2.2) (p, ) TRM rad Yaw commaad path i n t e g r a t o r output

3 55 C R c m

(figure 2.4) (r )

=TRM 56 C RINT rad I n t e g r a l of body axis yaw rete

( f i g u r e 2.4) (ruT)

Yaw forward path i n t e g r a t o r output 3 57 C RETRM rad

(figure 2 . 4 ) (re

TRM

C YTRM rad Yaw trim input ( f i g u r e 2.4) (rm)

3 58

3 C rad Pitch command p a t h i n t e g r a t o r output

QCTRM

( f i g u r e 2.3) (qcTRM)

A-7

Coded

blt8 Dtf i n i t i o n

- Segment Element Type Name

rad P i t c h forward path i n t e g r a t o r o u t - 3 6 1 C

2 . 3 ) (q )

put (figure

%ai

r a d P i t c h trim input (figure 2 . 3 ) (qm)

3 62 C Heave forward parh i n t e g r a t o r output

lb

3 63 C 2.5) (Te ) (figure

TRM

rad P i t c h hg1.e ( 8 ) 3 64 C rad 3 65 C 3 66 ' C rad

-

3 67 C Output of p i t c h 'nput shaping f i l t e r . .

( ( 1 ; )

-

3 68 C Output of r o l l Input shaping f i l t e r (P$

-

3 69 C Output of yaw input shaping f i l t e r

(r:)

Segment 4 - A i r c r a f t Mass, Geomtry and h e r t i a Constants

A.5

Coded

Segment E l e m M t Type Name Units De f in1 t i o n

Aircraf t weight (W)

1 ' N WT lb

N Ix s l u g f t

4 2 A i r c r a f t r o l l body a x i s i n e r t i a (Ix)

4 N IY s l u g f t A i r c r a f t p i t c h body axir: inei,r!.a (Iy)

A i r c r a f t yaw body a x i s i n e r t i a (IL)

4 4 N 12 s l u g E t

N s l u g f t A i r c r a f t cross product of inertia (I )

4 5 Ixz

XZ

* N FSCG f t Fuselage s t a t i o n of reference c g (FSCC)

4 6 N f t . B u t t l i n e of reference cg (BLCG)

4 7 BLCG

Uaterline of reference cg (acc)

4 8 N * m C G f t .

N f t / e e c Acceleratfon due t o gravity (8) 4 9 CC

-

EAC h e r t i , r a t i o = (I2 - IY)a ' X

4 10 C

-_

4 1 1 C PAC Inertia r a t i o = (IY - I X ) / ' Z

L

h e r t i a --+io - (12 - I x ) / I Y

4 12 C GAC

--

Croee Product of i;it?ttia r a t i o -

4 13 C JX

- m / r x

A-8

c d c d

N a W unit. Dcf inition

Type Segmeat Eleaent

+ 14 C J Y L Cross Product of inertia ratio =

- m/n

4 15 c JX Crore Product of inertia r a t i o =

- rla/Iz

4; 18 c M S s l u g A i r c r a f t mss (m)

4 1 7 c JlriuXJz -- 1 - JX*JZ

A . 6 F e m t 5- Tolerances f o r Errors in Trim Solutions

Coded Segment Eleaw. t Typ e Malle Units Definition

5 1-19 N * ( I ) Varies Specifies t h e acceptable t o l e i a c e

on t h e trin solution of the I t h

trim equation. The trim equatioas are given in Subroutine ERROR.

A-7 Segment 6 - Aero L i ' t and Drag Coastants

3ed SegmMt Elemeat Tpp B Nane Units Definition

1 N M a b -- Specifies whether aero data are to

be calculated (AkER) = 0) or are

stored i n t a b l e f o m t (AA€B# # 0 ) .

Note: For a l l scbscripted constants in Segmznt 6 ; I = 1 correspo I = 2 t? horizontal s t a b i l i z e r (includes foruc_,- t o wing data,

I = 3 t o vertical

canard-aod.aft-cmventional-locations), s t a b i l her.

?,25,48 GeoPretric or reference aspect r a t i o of I t h l i f t i n g surface (AR)

3,26,49 Aspect r a t i o of exposed area O F I t h

L i f t i n g surface (ARe) 4,27,50 Geoutric o r reference area of I t h

l i f t i n g surface <S)

5,28,51 Exposed area of I c h l i f t i n g surface

is,)

6,29,52 AqR#(5,I) rad Sueep of geometric or reference

quarter chord of Ith l i f t i n g s t r f a c e /*) 7,30,53 AEM(6,I) rad Sweep of leading edge of I t h l i f t i n g

surface a , )

Reference t a p e i r a t . ' o of I t h l i f t i n g

8,31,54 P:R@(,,x) -

sdrface ( h )

A-9

Coded

Scgmzn t Element Type Name units Dtf i n i t i o n

BfaxLmm lift c o e f f i c i e u t of I t h

6 9.32.55 I

l i f t i n g murface <%)

&le of a t t a c k f o r me.xI.ur l i f t

6 10.33.56 N

c o e f f i c i m t of I t h l i f t i n g s u r f a c e

(%)

N Sectional l i f t curve slope of Xth

6 11,34,57

l i f t s u r f a c e (a o)

Zero l i f t drag c o e f f i c i e n t of It!

6 12.35. -8 N

l i f t i n g s u r f a c e (CDo)

Ratio of fuaelage width a t l i f t i n g

6 13.36.59 N

surface i n t e r s e c t i o n to l i f t i n g

surface span for I t h l i f t i a g surface

Cd/M chaage in l i f t c o e f f i c i e n t a t zero

6 N

1 4 . 3 1 . 6 0

angle of a t t a c k per u n i t t r a i l i n g edge f l a p & f l e c t i o n :or I t h l i f t i n g surface

( ACb/bTEP)

N Chsnge in l i f t c o e f f i c i e n t a t angle

6 15,38,61 h E R $ ( 1 4 . 1 ) l/rad

of a t t a c k f o r m a x i m l i f t per u n i t trailing edge f l a p d e f l e z t i o n f o r I t h l i f t i n g surface.

( A c 4 u r x i l , F )

N A&( 15. I) 1/ rad Change in drag c o e f f i c i e n t per u n i t

6 16.39.62 t r a i l i n g edge f l a p d e f l e c t i o u f o r

~ t h l i f t i n g surface (4~,,/ & , $

Change in ritching PPoPent c o e f f i c i e n t

6 17.40.63 I &(16.1) l / r d

per u n i t t r a i l i n g edge f l a p d e f l e c t i o n

f o r I t h l i f t i n g surface i A C m / ( T E F )

I Change in angle of a t t a c k f o r maximum

6 18.41.66 l i f t per unit leading edge f l a p d e f l e c t i o n f o r I t h l i f t i n g s u r f a c e

( A 4 wldmF)

6 19.42.65 N Change in l i f t c o e f f i c i e n t a t zerd

angle of a t t a c k p e r i m i t leading edge f l a p d e f l e c t i o n f o r I t h l i f t i n g

N AeR@(l9, I) l / r a d Change i n l i f t c o e f f i c i e n t a t angle

6 20.43.66 of attack f o r maximum l i f t per u n i t leading edge f l a p d e f l e c t i o n f o r L t h Coded

Segment Element Type Name hits Def Inition

6 21,44,67 m ( 2 0 . I ) l l r a d Change in drag c o e f f i c i e a t per unit

leading edge f l a p d e f l e c t i o n f o r

I t h l i f t i n g s u r f a c e (ACi,/a ,&

i . 22,45,68

& ( 2 1 , 1 ) l / r a d change in pitching moment c o e f f i c i e n t

.

per unit leading edge f l a p d e f l e c t i o n f o r I t h l i f t i n g s u r f a c e ( ACp/( LEp) 6 23,46,69 AEB#(22,1! t a d Difference betveen angle of a t t a c k f o r zero leading edge f l a p e f f e c t i v e n e s s and angle of a t t a c k f o r maximum l i f t ,

i.e. when (ALPA-GE. (AERO(22,I) +

AEIu)(9,I)) t h e leading edge f l a p e f f e c t i v e n e s s equals zero (W =)

& ( 2 3 , f ) rad Difference between angle of a t t a c k

6 24,47,70 N

foz zero trailing edge f l a p e f f e c t i v e - nes8 and angle of a t t a c k f o r maximum l i f t , i.e. when (ALFA.=. (AERO(23,I)

+ AERO(9,I)) t h e t r a i l i n g edge f l a p

e f f e c t i v e n e s s equals zero (bd

TEF)

N SURFs(1,I) - Noraalized ( t o I t h l i f t i n g s u r f a c e

6 71,74,77 rean aerodynamlc chord) const m t €or c a l c u l a t i n g center of pressure s h i f t of f t h l i f t i n g s u r f a c e as a function of angle of a t t a c k , Appears as m u l t i p l i e r of cosine terms i n t h e equations (CP,)

N -- Normalized ( t o I t h l i f t i n g s u r f a c e

6 72,75,78 SuRFcP(2,I)

wan aerodynadc chord) constant f o r

c a l c u l a t i n g center of pressure s h i f t of I t h l i f t i n g surface as a function Appears as of angle of attack.

m u l t i p l i e r of s i n e terms i n t h e

equations (e2 j

N SURPCP(3,I) -- Normalized ( t o I t h l i f t i n g surface

6 73,76,79 m a n .aerodynamic chord) constant f o r c d c u l a t i n g center of pressure s h i f t of I t h l i f t i n g s u r f a c e as a Appears function of angle of a t t a c k .

only in the high angle of a t t a c k equation f o r center of pressure s h i f t 6 80-82 C CNMo(I) l / r a d Constant ca! culated by Subroutine

FORCES f o r use by Subroutine LIFDRG

t o generate nonlinear I t h l i f t i n g

s u r f a c e l i f t coefficient(C 1

N*O A- 1 . 1 Coded

S e m n t Element Type Nane lkli La Definition

6 83-85 C A g m ( 1 ) l/rad Conetent calculated by Subroutine

PORCES f o r we by S u b r o u t h e LIFDRG

t o generate nonlinear I t h l i f t i n g ( E ) .

surface l i f t c o e f f i c i e n t .

6 86-88 c * S ( I ) - Osvald efficiency f a c t o r f o r Ith

l i f t i n g surface calculated in Subroutine popI%s (e)

6 89-91 c CIA(1) l / r a d -ear l i f t curve slope of Ith

l i f t i n g surface calculated i n Sub-

routine FORCES (% )

* e

Se-t 7 - Wing k r o Constmta

A.8 Coded

segment Element TpPe Nare u n i t s Definition

7 1 * N Fsw f t Fuselage s t a t i o n of ving r e f t r e n c e

center of pressure (FS 1

7 2 N B M f t B u t t l i n e of reference center of

pressure of l e f t wing ( B \ ( l ) )

7 3 N BLM2 f t Buttline of reference center of

pressure of r i g h t wing ( B b ( 2 ) )

7 4 N W f t Waterline of wing reference center

of pressure (id+,) N C W f t Mean aerodynamic chord of wing (e,) 7 5

7 6 N .BU f t Reference wing span (b )

W

7 7 N E K D A l - Change in l e f t ving anple of a t t a c k

per u n i t deflection of i e f t wing control surface (Kb (1): C

7 8 N E K M 2 - Change in r i g h t wing angle of a t t a c k

p e r unit deflection of r i g h t wing

(K 4 (2))

c o a t r o l surface C

N EIU rad Wing Incidence angle

7 9

7 10 N CLBW

of wing (CI 1

Q O

7 11 l / r a d Change i n wing C A per u n i t >f

N CLBCLW

ving l i f t coeffic e n t ( C 1 f

%

A-12 Coded

m ; t o Definition

Segment Element N -

Type

12 N change ia wing C t , per u n i t of

7 l/rd

vlng lift coefficient ( C a 1 :

CL'

7 13 N 11 rad

7 14 l / r a d N per u n i t of

7 15 N l l r a d Change in wing Cn

ving l i f t c o e f f i c k t squared.

Change i n wing C& p e r u n i t of wing 7 16 N l l r a d

l i f t coeffjcienc squared (C, 2)

fCL

per u n i t of wing 7 17 N l l r a d Change in wing C,

drag c o e f f i c i e n t 1% )

' C D

7 18 N l / r a d Change in wing kP p e r u n i t of wing

l i f t c o e f f i c i e n t (& )

pCL

-

Not used

7 19 N

u

7 N Zero l i f t pitching mmrnt c o e f f i c i e n t

of wing (\>

-

Nct used

7 21 N

--

Nunber of e n t r i e s in elevon

7 22 N

effcctlveness t a b l e ( 4 7) p e r unit of wing

23 N l l r a d Change in wing CY

l i f t c o e f f i c i e n t !quared (CyB 2) CL

X h d y ui. disitance between a i r c r a f t

7 24 C f t

reference cg and left wing reference

center of pressure (kNG(l))

7 25 C f t X body a x i s distance between a i r c r a f t

reference cg and r i g h t wing iefe''ence

center of pressure (SmG( 2) )

Y body axis distance between aircraft

7 26 C f t reference cq and l e f t wing reference

center of pressure (lmNG(l.))

A-13 Coded Units Def Inition Segment Element Type N a m e

7 27 c Y a N G 2 ft P body axis d i s t a u c e between a i r c r a f t

reference cg and r i g h t wing r e f e r a c e

center of pressure (YmG(2))

C ZWINGl f t Z body axis distance between aircraft

7 . * 28 reference cg and l e f t wing reference center of pressure (hG(l))

7 29 c mING2 ft Z body axis distance between a i r c r a f t

reference cg and r i g h t wing reference

center of pressure ( k N G ( 2 ) )

Segment 8 - W r i z o a t a l S t a b i l i z e r Aero Constants

A . 9

Coded

Seg.eat Element Type N a a e unit 8 Definition

f t Fuselage s t a t i o n of horizontal 8 1 s t a b i l i z e r reference center of

pressure (Pk)

8 2 f t B u t t l i n e o f horizontal s t a b i l i z e r

reference center of pressure (BLfl) f t Waterline of horizontal s t a b i l i z e r 8 3

reference center of pressure (%)

8 4 f t Mean aerodynamic chord of horizontal

s t a b i l i z e r (g)

8 f t Reference span of horizontal

s t a b i l i z e r (h)

-

8 6 Change i n h a r i z o n t a l s t a b i l i z e r angle of a t t a c k per u n i t deflection of

ho r izon tal s t a b i l i z e r con t ro 1

surface (kd )

e

8 7 C o r N EIH rad Horizontal s t a b i l i z e r incidence ankle

(h)

8 rad Change i n downwash =-.le p e r u n i t

of wing l i f t c o e f f i c i e n t (e/cLw)

-

9 Not used

_-

8 10 Zero l i f t pitching mment c o e f f i c i e n t

of horizontal s t a b i l i z e r (f ) -€I e 8 11 Not used A 4 4 Coded

Segment Element Type Name units Definition

8 12 C X H f t X body axis distance between a i r c r a f t

cg and horizontal s t a b i l i z e r ref-

erence center of pressure (f)

13 c m f t Y body axis distauce between a i r c r a f t

8 .

cg and h o r i z o n t a l s t a b i l i z e r reference center of pressure cu,,

8 14 C ZH f t 2 body a x i s distance b e t w e e n a i r c r a f t

cg and h o r i z o n t a l s t a b i l i z e r reference

center of pressure (zK>

8 15 c XWeL f t ’ Constants r e l a t e d t o r e l a t i v e wing-

8 16 c m L f t h o r i z o n t a l s t a b i l i z e r location and

8 17 C AUEL rad which are w e d t o c a l c u l a t e dynamic

8 18 C D U U - pressure a t t h e horizontal s t a b i l i z e r

%EL* %EL’ %EL* %EL’

A.10 Segment 9 - Vertical S t a b i l i z e r Aero Constants

Coded

Segmen’ Elemeat Type Name units Definition

9 1 N FSV f t Fuselage s t a t i o n of v e r t i c a l s t a b i l i z e r

reference center of pressure (F%) 9 2 N

BLV f t B u t t l i n e of v e r t i c a l s t a b i l i z e r

reference center of pressure (Bk,)

9 3 N

WLV f t K a t s l i n e of v e r t i c a l s t a b i l i z e r

refezence center of pressure (Wq,)

9 4 N Mear aerodynamic chord of v e r t i c a l

cv f t

s t a b i l i z e r (cv) 9 5 N BV f t Reference span of v e r t i c a l s t a b i l i z e r (bv)

E W R u

9 6 N Change i n v e r t i c a l s t a b i l i z e r angle of a t t a c k p e r unit d e f l e c t i o n of v e r t i c a l s t a b i l i z e r c o n t r o l surface

( K 4 6

r

7 C o r N KIV rad V e r t i c a l s t a b i l i z e r incI4ence angle (i,)

--

8 N EKJI Sidewash c o e f f i c i e n t (Kv)

--

9 N CNOV Zero sice farce yawing moment co-

e f f i c i e n t of v e r t i c a l s t a b i l i z e r (Cn ) V Coded Segment Element Typc Name unitr Definition

10 N A g l I W D -

N u d e r of e n t r i e s in rudder effective-

ness t a b l e (S 7)

Angles of a t t a c k f o r e n t i r e s i n

9 11-20 N TAtFA(1) rad

vertical s t a b i l i z e r effectiveness

t a b l e (1-1 t o AALFA)

-

Number of entries in vertical

9 21 N W A

s t a b i l i z e r e f f e c t i v a e s s t a b l e ( i. 10)

Vertical s t a b i l i z e r effectiveness

9 22-31 N TVTKFF'CI) -

t a b l e e n t r i e s (I = 1 t o AALFA)

(f3mRo (4 1) . -

9 32 c x v f t X body axis distance between aircraft

cg and vertical s t a b i l i z e r reference

center of pressure (5)

9 33 c w f t Y body axis distance between a l r c r a f t

cg and vertical s t a b i l i z e r reference center of pressure (yv) Z body a x i s distance betveen a i r c r a f t

9 34 c zv f t

cg and vertical s t a b i l i z e r reference center of pressure (+)

Segment 10 - Puselage Aero Coastants

A . 1 1 Coded

Segment E1cllrPm.t Type Name b i t 8 Definition

10 1 N FSF f t Fuselage s t a t i o n of fuselage

reference center of pressure (FSF) Buttline of fuselage reference 10 2 N 'BLF f t center of pressure (BLF)

10 3 N W L F f t Waterline of fuselage reference

(%)

center of pressure Maxirim fuselage a x i a l force normal-

10 4 N A d Q O ft2

ized t o dynamic pressure( (A/q)pWO

10 5 N S h K n f t 2 Maximum fuselage s i d e force normalized

t o dynamic pressure( (S/q)=)

10 6 N N m M f t 2 Maximum fuselage normal force normalized

t o dynamic pressure ((t+iF/q)MAx)

10 7 N l#Q?fXl f t 3 Maximum fuselage p i t c h moment in

AlF -

region ( 'Tf / I 8 + AOF) 4 aS 6

normalized t o dynamic pressure A-1 6 Coded

Segment Element Type Name wits Definition

Haximum fuselage p i t c h moment i n

10 8 N

- norulalized

region - A l F 4 o ( gfl

t o dynamic pressure ( ( W q ) m 2 )

N Maximum fuselage yaw moment i n region -

10 9

NbQMxl

% / l a f .<% f A2F - n o m a l i z e d t o

dynamic pressure where 4~ fuselage s i d e s l i p angle ( (8/q)ml)

10 10 N Haximum fuselage yaw moment i n region -

Nfim

A2F.L 4 ~ ~ 4 % - normalized t o

dynamic pressure ((N/q) MAX2 Angle of a t t a c k where fuselage normal 10 11 A O F f o r c e equals zero ( r4 0 , ) Angle of a t t a c k where fuselage p i t c h

10 1 2 AlF

moment equals zero (d )

IF

Fuselage s i d e s l i p angle where fuselage

10 13 A2F

yaw moment equals zero (d ) 2F Exponent of cosine shaping function 10 1 4 EN1 f o r fuselage axial f o r c e (nl) Exponent of sine shaping function f o r EN2 10 15

fuselage side force ( n , )

Exponent of sine shaping function for

10 16 EN3

fdselage normal force (nj) EN4 Exponent of s i n e shaping function f o r 10 1 7 in t h e region fuselage p i t c h moment

( n/18 + AOF) L, d A1F (n4)

of s i n e shaping function for Exponent

10 18 N EN5

fuselage p i t c h moment i n t h e region A l F e d S ' l f ( n , ) Exponent of s h e shaping function f o r

10 19 N EN6

fuselage yaw moment i n t h e region

n / l 8 f d F~ g A2F (n,)

10 20 EN 7 Exgonent of s i n e shaping function f o r

fuselage yaw mment in the region

A2F fr as Fy s n (n7)

Maximum a f t s h i f t of fuselage c e n t e r of

10 21 DFSW pressure with angle of a t t a c k ( b F S ) *O A-17 Coded

Segment Elenen t "me Name Units Def i n i t i o n

10 22 N ADFSFl rad Angle of a t t a c k where fuselage center

of pressure begins t o s h i f t a f t

10 23 N A D F S n rad Angle of a t t a c k where fuselage center

of pressure s h i f t equals DFSFO (1LpSz).

-

10 24 N XKYF Constant t o control fuselege s i d e

force contribution t o fuselage yaw moment (KyF)

10 25 C XF f t X body axis distance between aircraft

cg and fuselage reference center of (XF) pressure

YF f t Y body axis distance between a i r c r a f t

10 26 C

cg and fuselage reference center of

pressure (YF)

Z body axis distance between aircraft

10 27 C ZF f t

cg and fuselage reference center of pressure (Z,)

Segment 1 1 - Direct Thrust Force and Moment Constants

A . 1 2 Coded

Segment Element Type Name Units Definition

f t Fuselage s t a t i o n of engine nozzle

11 1-2 N FSSW(1)

swivel point; I - 1 f o r l e f t engine, = 2 f o r r i g h t engine (FSsw(1)) Buttline of engine nozzle swivel

11 3-4 N BLSW(1) f t

point; I = 1 f o r l e f t engine, = 2 f o r r i g h t engine ( B L w ( I ) ) Waterline of engine nozzle swivel

5 - 6 N f t

11, WLsW(1)

point; I = 1 for l e f t engine, 5 2 f o r r i g h t engine (WLsw(1)) SIGY rad I n c l i n a t i o n of engine c e n t e r l i n e t o

11 7 N

fuselage r e f e r e r c e l i n e ( W Y )

Length of engine nozzle (hoz)

11 li N E U f t

Change of engine nozzle t h r u s t

11 9 N KAFTl l l r a d

e f f i c i e n c y per unit t h r u s t turning angle (y.AFT1) Coded

Segment Element Type Name mite Definition

11 10 N W T 2 L Engine nozzle t h r u s t e f f i c i e n c y

a t zero t h r u s t turning angle (KmT )

11 11 N -- Number of engines ( 6 2 ) (n )

AENG m g

1 1 12 C CSIGY _- cos ( C y )

1 1 13 c

XT (I) ft X body axis dtstance between

1 1 14-15 c

aircraft cg and nozzle swivel point;

I = 1 . f o r l e f t engine, = 2 f o r r i g h t

engine ( % ( I ) ) .

YT (1) f t Y body axis distance between a i r c r a f t

11 16-17 C

swivel point; I = 1 f o r

cg and nozzLe

left engine, = 2 f o r r i g h t engine

(YT (1 11

ZT(1) f t 2 body axis distance between aircraft

11 18-19 C

cg and nozzle swivel point; I = 1 f o r = 2 f o r r i g h t engine l e f t engine,

( % ( I ) 1

A.13 Segment 12 - hal.at Ram Force and Moment Constants

Coded Segment _Element Type i 4 - Units Definition Fuselage s t a t i o n of engine i n l e t force; 1-2 N f t

I = 1 f o r l e f t engine, = 2 f o r r i g h t

engine (FSZN(I) 1

f t Buttline of engine c e n t e r l i n e a t

12 3-4 N

inlet face; I = 1 f o r l e f t eqgine,

- 2 f o r r i g h t engine (BLIi(I))

Waterline of engine c e n t e r l i n e a t 12 5-6 N f t i n l e t face; I = 1 for l e f t engine

= 2 f o r r i g h t engire (ULIN(I))

f t Equivalent inlet radius ( r m )

12 7 N

--

Number of e n t r i e s i n V/Y, t a b l e ( 4 5)

12 8 N

-- Number of e n t r i e s i n ram moment am,

12 9 N

ram e f f i c i e n c y and incremem *-a1turning

angle t a b l e s ( 6 8 ) A-19 Coded

Segment Element- T y p e Name Unite D e f i n i t i o n

12 10-11 c f t X body axis d i s t a n c e between a i r c r a f t

cg and wine i n l e t ; I = 1 f o r l e f t

2 f o r r i g h t engine ( % ( I ) )

engine,

12 12-13 c mu) f t Y body axis d i s t a n c e between a i r c r a f t

cg and engine i n l e t ; I = 1 f o r l e f t engine, = 2 f o r r i g h t eugine (Ym(I))

12 14-15 c ZIN(1) f t Z body axis d i s t a n c e between a i r c r a f t

cg and engine i n l e t ; I = 1 f o r l e f t

engine, = 2 f o r r i g h t engine (Zm(1)).

Segment 13 - Coriolis Force and Moment Constants

A.14 Coded

Segment ElelPent T y p e Name Units Definition

13 1 N ELDUCT f t Length of engine duct (lDvcT)

Segment 14 - Propuleion System Constants

A.15 Coded

Segment Element T y p e Name Units Definition

1 4 1 AMACH -- Number of e n t r i e s in Mach number

t a b l e ( 4 4)

KBTl Fractional change in t h r u s t p e r

1 4 2 1 /%

percent bleed (K )

BT1

Fractional t h r u s t a t zer:) bleed @aT2)

1 4 3 KBT2

K M l - Fractional change In I n l e t mass flow

1 4 4

r a t e p e r f r a c t i o n a l change In t h r u s t

(b$

- Fractional inlet mass flow rate at

KM2 1 4 5

zero f r a c t i o w l t h r u s t (K )

m2 rad/sec Haximum engine r o t a t i o n a l speed in 14 6 pME=

rad/eec RM)

14 7 RPHMX I h i m u m engine r o t a t i o n a l speed in rPm revolutious per minute 1 4 8 RCS L Indicates whether reaction control

s y s t e m IR a c t i v e ; RCS # 0, RCS is

a c t i v e ; 0, RCS not a c t i v e A-20

Co&d

Sement Elcaurrt TYPC loam Ualtr Dcf i n i t l o n

Number of entries in f r a c t i o n a l t h r u s t

14 9 N

t a b l e f o r determining engine r o t a t i o n a l

speed ( 4 6 )

14 10 N JENG e l t q f t Engine r o t a t i o n a l inertia per engiue

(JENG)

AFRAC --

lumber of entries i n f r a c t i o n a l t h r u s t

1 4 1 1 N

t a b l e f o r engine dynamics parameters

( f 10)

T A W sec T h e constant f o r afterburner t h r u s t

14 12 N

dynamics ( T A B . > Acceleration l i m i t of afterburner

14 13 N ABACCEL llsec

thrua t (T AB&

1 4 14 N AEECEL Usee Deceleration l i m i t of afterburner

t h r u s t (T*N)

--

N

14 15

TABdN +w+l of TABIN when afterburner l a

lit (T ) ABON

TABOFF -

14 16 N Level of TAB(1) when s f terburner is

turned off .TAB 1

OFF

C Fractional t h r u s t level a t reference 1 4 17 bleed (KBT)

h.16 Segment 15 - RCS Force and Moment Constants

Coded Segment Element Type N a ~ e Unics Definition

15 1 N ARCSJET -- idumber of RCS jets ( f 10)

15 2 N ABLDR - Number of entries i n bleed tables (+6)

15 3 N ELDREX x Reference bleed (BmF>

15 4-13 N FRCSMX(1) !b Maximum force of I t h j e t (FRcsMx(I))

15 14-23 F DEMAND(1) -- Indicates whether j e t I is demand

bleed; DEMAND(1) = 1, jet I is demand;

= 0, jet I is continuous bleed ( % ( I ) )

15 24-33 N BLDM~R(I) -- Indicates whether jet I requests

a d d i t i o n a l bleed; BLDMOR(I)=l, j e t I

requests additional bleed; - 0, jet

I&t operate with a v a i l s b l e bleed

(BLDM ’ 1) )

A-2 1 Coded

S a m . - Eleeent Type Name unit8 De f i n i t Ion

-

Daftaarr whether J t h control drives

N BCSL(1.J) -

15 34-83

I t h jet; RCSL(1,J) # 0, J t h control

drive8 fth jet; - 0 , J t h control does

not drive I t h jet. J = 1 r o l l

control; J = 2, p i t c h control; J = 3

yav control; J - 4, s l d e force control; J = 5, normal force coatrol Fuselage s t a t i o n of I t k jet (FSJET(I):

N

15 84-93 Buttline of I t h j e t (BLJET(I))

N

15 94-103

N Waterline of f t h jet (WLqT(T))

104-113 Pitch o r i e n t a t i o n of jet chrust line

N

15 114-123 (O&T(') )

N PSIRCS(1) rad Yaw o r i e n t a t i o n o f j e t :.hrus' line

15 l24-133

(Y',(I) 1

134 C FRCSMKR - Not used

FRCSMXO l b Total RCS force produced a t reference

15 135 C

conditions (reference bleed and zero control inputs); i f a l l jets are demand bleed, then FRCSMXO = 0 ( F R C S q , ) 136 C FRcsMxl -- Not uec. .

C'IIITRCS(1) -

15 137-146 C 'OB ( e x T ( l )

STKTRCS(1) --

15 147-156 C sia (eaT(I))

157-166 C CPSIRCS(1) -

15 COB (qJET(I))

SPSIkcs(L) -

167-176 C sin ( P J e T ( I ) : C WrT(1) f t X body axis distance between a i r c r a f t 15 1 7 7-186

cg and I t h jet (XJdTfI))

Y body b x i r p distance between a i r c r a f t 187-196 C YJET(1) f t 'g and Ith jet (YJET(l))

ZJET(X) f t Z body i s distance between a i r c r a f t

197-206 C

cg and ~ c h jet (ZJET(I))

A-22

A.17 &went 16 - Control System Constants

CO&d

" e m n t Eement lbPe H8me Units Definition

-

16 I 60t used

-

16 2 I l o t used in f i n a l version of program.

-

iC N Hot used i n final version of program.

-

16 A B o t used in f i n a l version of program.

-

16 A Hot used in f i n a l version of program.

-

1; I A Bot used i n f i n a l version of program.

16 B see pilot gain on ZEDO!C feedback for

stationkeeping option

H SeC Pilot gain on YHwrr and XHWP feedback

1 0 8

i n lateral posicion loop for s t a t i o n - keeping option

16 N sec P i l o t gain on XEDOT and YEDOT feed-

back i n longitadha1 position loop f o r stat ionkeeping opt ion.

16 l l/sec ='?at heave rate c o n t r o l l e r gain f o r

stationkeeping option lateral s t i c k gaPn f o r s t a t i o n -

16 1 1 I rad/ft P i l o t

keeping option

16 12 N rad/ft P i l o t longitudinal stick gain f o r

stationkeeping option

-

Hot used i n f i n a l version of program.

16 N

16 N f't/sec Control system function switch speed

(v 1

*SW

-

N Not used i n f i n a l version of program.

16 15

-

H lot used i n f i n a l version of program.

16 16

19 rad Roll control system input gain table

16 17-24

if the gain is a function of p i t c h angle.

B rad P i t c h angle table for determining

16 25-32

control system gains.

14 l / r a d Roll control s:rstem e r r o r gain table if

16 33-40

t h e gain is a f w x t i o n of p i t c h angle.

16 41-48 N rad/sec Roll control system input i n t e g r a t o r

gain table i f t h e gain is a function of p i t c h angle.

A-23

c-L

Coded

Segment Elemcn t Type N a m h i t 8 Def l n i t l o n

-

Y u b . r of entriem in c o n t r o l system

49 Y

pip. p i t c h angle t a b l e ( Z 8)

Y rad/utc/ lbll control sya tar error l i t t e g r a t o r

16 50-57

rad &.in t a b i s +

Is r d r a d l Roll caotrol syatar roll rate

16 5 8 4 5 8 e C feedback gain table*

--

Output l i m i t on r o l l caatrol

Y 16 66

mystem error I n t e g r a t o r (p

eLm)

rad lbll control t o yav c o n t r o l inter-

16 67 I

caunect gain (IC4 6 )

Y r

P i t c h thrust d e f l e c t i o n to r o l l

N rad

16 68 c o n t r o l g& )

r r

rad D i f f e r e n t i a l w i n g tratllrrg edge

69 N

f l a p ( d a w n ) d e f l e c t i o n t o roll

C o n t r o l gain (Kd.1

t8d Yaw control system Input galn t a b l e *

16 70-77 N

Is 11 rad Y a w control system e r r o r gain table’

16 78-85 Y a u c o n t r o l s y s t e m input I n t e g r a t o r

86-93 N radlaec

gain t a b l e

N t a d / w c / Yaw c o n t r o l system error i n t e g r a t o r

16 94-101

rad gain t a b l e *

Yaw c o n t r o l ayzrtem yaw race feedback

N rad/ rad/

16 102-109

gain t a b l e *

sac

-

Output limit on yaw c o n t r o l s y s t e m

110 N

e r r o r I n t e g r a t o r (t eL1d

iia U rad/8ec Yew c o n t r o l system trimer gain

112 N :ad Y a u t h r u s t d e f l e c t i o n t o yaw control

gain m y * 5 y)

Rudder d e f l e c t i o n t o yaw c o n t r o l

113 N Lad

C q , )

Lateral acceleration feedback gain

N r a d / f t /

16 114-121 S@C2 tabLe; loaded as a function of T Q W ( I ) A-24

16 122 I Thnut colQd t o yaw control gain % )

Y

16 I23430 I

P i t c h coatrol system input gain t a b l e *

16 131-138 N

Pitch control system error gain, table*

I Pitch control system input integration

16 139-146

tabl&

P i t c h control aye- error i n t e g r a t o r

16 1470- I rad/.=/

rad gain table"

P i t c h control system p i t c h rate feed-

16 US-162 U rad/rd/

back g a b table"

sac

c

Output lirit ou p i t c h control system

16 163 I

error iategrator ( ~ ~ m )

I rad/sec P i t c h control ayefa tri-r gain

16 164 I W h e n the absolute value of longitud-

165 N

inal s t i c k d e f l e c t i o n I s greater than t?SllC, automatic p i t c h t r i m I s a c t i v a t e d wing t r a i l i n g edge f l a p

166 N rad S ! t r i c a l

(elevon) d e f l e c t i o n to p i t c h c o n t r o l g a b P i t c h t h r u s t d e f l e c t i o n t o p i t c h

16 16 7 N rad

control gain

16 168-1 75 N l b / f t / m c Heave control system e r r o r gain table*

16 176-183 N lXZBIl(1) l b / s e c / Heave c o n t r o l system error I n t e g r a t o r

f t / s e c gain table"

184 I l b Thrust c o r e a d t o manual t h r o t t l e

input

Z B U l b Output l i m i t on heave c o n t r o l system

16 185 N

error i n t e g r a t o r (Tern)

N b f e r e r r c e elevon d e f l e c t i o n ( 6mF EWINR? r a d

16 186

%

PSITILF(1) rad Referaace yaw t h r u s t d e f l e c t i o n ;

16 187-180 N

I - 1 f o r l e f c engine, =2 f o r r i g h t

engine ( v ~ ~ ~ ( 1 ) )

KEFLAPl - Change in h o r i t o n t a l s t a b i l i z e r

189 N

trailiag adge f l a p d e f l e c t i o n p e r mlt

change In angle of a t t a c k (Kmm)

'Provided +,,he gain is a function of A-23 pitch angle.

coded

Sep.aa;t Elemlot T y p . r i a n %ita Definition

16 P

16 191 I m e in w i n g leading e - f l a p

d e f l e c t i o n per unit change in

angle of a t t a c k (%)

16 192 N Uing leuding edge f l a p deflection

at ram angle of a t t a c k

U

16 193 U & r of entries in dynamic pressure

table for lateral rccaleratiar

feedback gain ( 4 8)

16 194-201 I Dynamic pressure t.?ble f o r lateral

acceleration feedback gain

I Pi+,ch input shaping filter t h e

202-209

constant table if the t i m e constant

is a ftmction of p i t c h angle R o l l input shaping filter time constant

H

16 210-217 table if the time constant is a function of pitch angle

I Yau i n p u t shaping filter t i m e ccnstant

218-225

t a b l e i f t h e t i m e constant is a m c t l o r

of p i t c h angle

A.18 Segrant 17 - Actuation System Conatants

Coded

Segwmt Bl.emmt Type N m hit& Definltioa

k f t wFag t r a i l i n g edge f l a p (elmmm)

N radlsec

actuator constants in order from

rad/sec

17 2 N

elererrt 1 t o elercot 5 : upper rate

Is rod

1 7 3

rad l i m i t , lower rate lipit, 'pper

I

17 4 8 e C pooitioa lirit, lover position l i m i t ,

N

17 5

time constant

N radlsec Right w i a t r a i l i n g edge f l a p (elevon)

actuator coastante in order from

N tadlaec

dement 6 t o cla-t 10: uppor rate

N rad

1 7 8

l i m i t , lover rate limit, upper position

N rad

lipit, lower position limit, time

I .(LC

17 10

constant

rad/rec Vertical s t a b i l i z e r control surface

1 1 N

1 7

radlaec (Rudder) actuator constuuto in order

12 N

from element 1 :

N rad t o eleirent 15: upper

rate linit, lover rate lWt, upper

I rad

poeiticm l i m i t , k e r position limit,

N eec

time constant A-26 Coded

Semen t Element TYW N - units De f i n i t ton

tbrltantal s t a b i l i z e r t r a i l i a g edge 1 7 16 3 HFURATL rad/aec

f l a p a c t u a t o r constauts in order f r o m

17 17 bl EFLRATL rad/.ec

N element 16 to elerent 20: upper rate

1 7 18 €IRAPMi' rad

17 I lirit, lower rate limit, upper

19 EFIAPX rad

i 7 20 N TIP setc position lirit, lower position liait, tire ccmstant

17 21 h' wR13uIz radlsec Uiag leading edge f l a p actuator

1 7 22 N WFLBATL rad/sec constaats in order from e l e k n t 2 1

1 7 23 t o el-t 25: u m r rate l i m i t ,

N rad

1ok-r rata l i m i t , upper position

1 7 24 N YPLApm rad

lidt, lower position l i m i t , time

1 7 25 N RIP 8 e C

c o n s t a t P i t c h thrust deflection actuator

N

1 7 26-27 -(I) rad/sec constants i n order (two elements-per

N

1 7 28-29 arnaL(1) rad/sec

constant where I = 1 is l e f t engine

17 I TElaUpL(1) rad

30-31

and I = 2 is r i g h t engine) from

1 7 32-33 N -(I) rad

element 26 to elePeat 35: upper

1 7 36-35 N lTm(1) sec

rate l i m i t , lover rate l i m i t , upper position l i m i t , lower position l i m i t , tioe constarit

1 7 36-37 ! a PSIWL(I) rad/sec Y a u t h r u s t d e f l e c t i o n actuator

1 7 38-39 N PSXLaL(1) radlsec constants in order (two elerPents p e r

constant where I = 1 is l e f t engine

1 7 40-41 N PSIWL(1) rad

N PSTI,PL(I! rad 8nd I = 2 is r i g h t engine) from

1 7 42-43 N el-t 36 t o aleuent 45:

1 7 44-45 TPSIT(1) sec upper

rate l i m i t , lover rate limit, upper

position l i m i t , lover position l i m i t , t- constant

N RCS area actuator constants in order

1 7 46-55 EGURL(1) --

N by groups of ten (up t o ten RCS j e t s

1 7 56-65 RCSLRL(1) -

can be specified) from d e m e n t 46 t o

1 7 66-75 N RCSUPL(1) --

17 76-85 N RcSLPL(1) - element 95: upper rate l i m i t (elements

1 7 N 'fAURCS(1) sec 46 to 55), lower rate l i m i t (elements

36-95 56 to 65) ,upper position l i m i t (elements 66 t o 75). lower position lidt

( e l e m t s 76 to 85), time conatants

(elemento 86 t o 95)

A-27 Coded S e m n t Element Type Name h i t e Def I n i t i o n

force d y n d c 8 CtmotPato In

1 7 96-105 U

order by group8 of tan (up t o ten

1 7 106-us v

R c s jet. c8Q be opeclfied) frcm

17 ll6-u5 I

el-t 96 t o el-t 145:

17 l26-135 19 upper rate

lirit ( e l m t r 96 t o 105), laver

1 7 136-145 I

rate l i m i t (elemnte 106 t o lU), upper pooition 1-t (elements 116

to US), later poaition lirit (dea~111t6

126 -to 135), tire constant (elereutlr 136 to 145)

146 N llrrirm c-ded symetrical wins

trail- edge f l a p (elevon) deflection ( &eL&

147 N MLM rad b x i u n c - d e d d i f f e r e n t i a l

wing t r a i l i n g edge f l a p (elevon)

deflection ( saLm)

148 I .DBLxM rad corended vertical s t a b i l i z e r

1 7 control surface (rudder) deflection

~ . 1 9 sagrmt 18 - Lift-Drag Variables

Coded S e m t El-t Type blam m i t a Definition .

Hiscel?xeous iLtermediate parameters

18 1 c L A m 1

18 2 c CIAB l / r a d associated with calcuiat* 'Q of

18 3 C B g W B & C N M O ( I ) , AER.OK(I), EOS(L), and

18 4 eel - - ( I ) . In V o l w I1 symbology, c c 2 -- these p a r a m t e r s are i n order: 18 5

-

18 6 c BJ

taahl/29 % ( B ) r cl* ' 2 and J

A.20 S e m t 19 - Uing Aero Variable0

Coded

S e m t El-t Type N - %its Def hition

19 1 c Dxw l b X body a d s wing aerodynamic force

( A++

19 2 C DYU l b Y body axis w i n g aerodynamic force

( AYw)

19 3 C DZW l b Z body axis w h - aerodynamic force

( b $ )

A-28

19 4 C ft lb

f t lb

1 9 s C

6 C ft lb

u

-

7 C

l9

Aver- local a u l a of attack of

19 8 C rad

-

mu1 ving coafflcirpt (

9 C

w

f t / n c X body ud. a i r @ at ving center

19 lo C

of ~ia..ura (ur,

ll C ftlwc Z body ui. ampeed at I V ~ center

of P - (*

I 2 C ft/.ac T o t a l ainpaed varrA by left ving

in m X-2 body .ri. plune at the

l a f t vine canter of presaure

(Va(1) 1 .

19 u C f t I 8 e C T o t a l airspaad oennd by rfght

kr an X-2 body a i m plane a t the

right canter o f presmra

( V X z ( Z ) )

Amraga t o t a l airspeed mmmd by

C f t / k C

19 14

l e f t apd right uing h l v e s (Vnr =

0 . 5 ( V n ( 1 ) + V & ) ) ) .

Change in mgla of attack of left

rad

19 u C

wiag due t o h f t wing-control eurface

&flection ( a q ql(1))

r d

19 16 C

r d

19 17

rad

19 18

--

19 19

w

l9 20

A-2 9

C0d.d

S.rraat n.-t m e - unit# Ikf inltlou

-

Ik.g co8fficiaot of l8ft

C

2 l

C % W

d

C Dog coefficieat of r u t a

(c4T(2) )

-

Mot wed

23 C

rut tiins a i d e d i p angle &)

2 1 C

l b / f t2 Dynamic premure based on Vxz(Q)

C

2s

QUS*BK

Wft

26 C

W C 0.5 BU/VXZ(b/V)

27 C

1s

Wing stability axis r o l l m m t ($ )

f t l b C S

f t Ib W i n g s t a b i l i t y axis yaw moment

C

l b Wing aide force (\ )

30 C

s

Left w 3 . q angle of a t t a c k (gw(l))

C rad L

sia (ALFl)

C

-

cos (ALFl)

C rad

C Right Wias angle of a t t a c k

34 (0(16(2))

-

Sia (ALPZ) 35 C

-

COS (u)

36 C

X body axis left wing aerodynamic

C l b force t&(l)) l b X b o d y b r i g h t wing aerodynamic 38 C force (S(2)) 2 body axis l e f t wing aerodynamic l b

39 C

force (z&))

l b 2 body axis right wing aerodynamic

40 C

force (+42))

Wing angle of a t t a c k ( a )

C rad

19 41

-

sin (ALF)

C

19 42

u

cos (ALP)

43 C

A-30 Coded.

SesPent Element Type N a m e bits Definition

urodyndc p i t c h aommt due

f t Ib

19 44 C

to flap8 and xero l i f t p i t c h (NB)

YIag body ui. roll m t without

1 9 45 C ft Ib

aaymetric l i f t e f f e c t . (h)

Uing body axis yav - t without

46 C f t lb

aapratric drag e f f e c t s (% )

B

k f t w i n g trailing edge f l a p

rad

1 9 47 C

d e f l e c t i o n (elevon) correctep f o r

l a r g e d e f l e c t i o n e f f e c t s ( 4 m m ( l ) )

Right w i n g trailing edge f l a g rad

48 C

d e f l e c t i o n (devon) corrected f o r

large d e f l e c t i o n e f f e c t s (6&42))

-

Large d e f l e c t i o n correction f a c t o r 19 49 C f o r l e f t wing elevon %%(1))

-

Lirge d e f l e c t i o n correction f a c t o r 19 SO C

f o r r i g h t wing elevon ( % F (2))

W

-

Left wing p i t c h -Peat c o e f f i c i e n t 5 1 C due t o f l a p s and zero l i f t p i t c h

(qqp) 1

-

Right Wiag p i t c h momat c o e f f i c i e n t 52 C due t o f l a p s and zero lift p i t c h

(cqJ2) 1

c Normalieed(to CU)dhift of left wing 19 53 C center of pressure as a function of

angle of a t t a c k ( b %INc(l))

-I

Normalized (to CW )&if t of r i g h t

19 54 C

wing center of pressure as a function

of angle of a t t a c k ( b ; ‘cIUc(2))

A . 2 1 Segment 20 - Horizontal S t a b i l i z e r Aero Variables

Coded .

Segment Element Wpe N a ~ v uni tr Def hit ion

20 1 C D X E P lb X body axis horizontal s t a b i l i z e r

aerodynaodc force (A%>

20 2 C DYE l b Y body axis horizontal s t a b i l i z e r aerodynamic force ( AYK: A-31

Coded

S e w a t Element TyFe N a m e Units De f i n i t ion

lb 2 body axis h o r i m n t a l s t a b i l i z e r

aerodynamtc f o r c e ( h % )

f t lb Horizontal stabilizer body a r i a

toll aoraat ( A$>

f t lb HorjZontal s t a b i l i z e r body axis

20 5

p i t c h moment ( A%)

f t l b b r i z o n t a l stabilizer body a x i s

20 6

paw moment ( A Ntl>

f t / s e c X body axis airspeed a t horizontal

20 7

s t a b i l i z e r center of pressure (I,@

f t l e e c

8 Z body axis airspeed a t horizontal

s t a b i l i z e r center of pressure (wH)

Aircraft angle of a t t a c k (a), a l s o

9 rad

coded as ALPH in segment 29

rad 20 10 1 1 rad Incremented angle of a t t a c k f o r hacking angle of a t t a c k rate e f f e c t s

( A a p r A 4 2)

20 12 c a1 Wing l i f t c o e f f i c i e n t a t DALFW where

increment is used

a m a M

(CL1)

_-

Wing l i f t c o e f f i c i e n t a t DALFW where

20 13 c cL2

a negative 4 increment is used (C$ _e

CD1 Wing drag c o e f f i c i e n t a t aALFw where

20 14 C a p o s i t i v e d increment is used (cD1)

--

Wing drag c o e f f i c i e n t a t DALFW where

20 15 c C D 2

a negative a ( increment is used

(C&) rad Change i n horizontal stabilizsrr angle

20 16 C DADE

of a t t a c k due t o h o r i z o n t a l s t a b i l i z e r

c o n t r o l eurface d e f l e c t i o n ( A U 6 )

e 'dorizontel s t a b i l i z e r l o c a l angle

20 17 ALFK rad

of a t t a c k (eE)

--

Horizontal s t a b i l i z e r l i f t c o e f f i c i e n t

20 18 CLB

(CQ) A-32 Coded

Segment Element Type Name units Def i n i t i o n

20 19 C Horhmtal s t a b i l i z e r drag co-

ef f icieat (

w

20 Variables related t o relative w i n g -

20 C

21 horixontal s t a b i l i z e r l o c a t i o n and

20 C

22 which are w e d t o calculate dyR-

20 C preeaure a t the s t a b i l i z e r ( A ,

20 23 C

U % .

%. Z) "

20 24 Normalized dynamic pressure l o e s at C

h o r i z o n t a l s t a b i l i z e r c e n t e r of

pressure ( A q/%)

Horizontal s t a b i l i z e r angle of a t t a c k

20 25 C

e tan-1 (=hi)) WIjl)

26 C sin (am)

COB (am)

27 C

28 C Change in h o r i z o n t a l s t a b i l i z e r

angle of a t t a c k due t o angle of

a t t a c k rate ( A d E )

20 C Horizontal s t a b i l i z e r control surface

( i . e . e l e v c t o r ) deflection (6,)

E f f e c t i v e dynamic pressure at 20 30 C

h o r i z o n t a l s t a b i l i z e r (&FF)

31 Average d e f l e c t i o n of l e f t and r i g h t 20 C

wing t r a i l j n g edge f l a p s ( &TEF*)

32 C Horizontal s t a b i l i z e r p i t c h mment

c o e f f i c i e n t due t o f l a p s (s)

33 Normalized (to CH) s h l f L : of h o r i z o n t a l

20 C s t a b i l i z e r center of prdttsure as a

function of angle .)f a t t a c k ( A ] k T )

h.22 Segment 21 - Vertical S t a b i l i z e r Aero Variables

Coded Segment Element Type N- U n i t 8 Definition

21 1 C D X V l b X body axis v e r t i c a l s t a b i l i z e r

aerodynamic force ( A Xv)

21 2 C D W l b Y body axis v e r t i c a l s t a b i l i z e r

aerodynamic force ( A Y , )

A-33 Coded

Segment Element Tvpe Name Unite Definition

Z body axir vertical s t a b i l i z e r

3 C DZV l b

aerodynamic force ( +)

f t lh Vertical s t a b i l i z e r 'body axis r o l l

21 4 C DLV

m-t ( U v )

f t l b Vertical s t a b i l i z e r body a x i s p i t c h

21 5 C DMV

=wme (A%)

f t l b Vertical s t a b i l i z e r body axis yaw

21 6 C DNV

m t (ANv)

W f t l s e c X body axis airspeed a t v e r t i c a l

21 7 C s t a b i l i z e r center of preesure (av)

f t l s e c Y body axis airspeed a t v e r t i c a l

8 C W s t a b i l i z e r c a t e r of pressure (vv) C DADR rad Change i n v e r t i c a l s t a b i l i z e r angle 21 9 of attack due t o v e r t i c n l s t a b i l i z e r

control surface d e f l e c t i o n ( Ad6 =)

Vertical s t a b i l i z e r l o c a l angle of 21 10 C ALFV rad a t t a c k ( 4 v)

--

Vertical s t a b l l i z e r l i f t c o e f f i c i e n t 1 1 C CLV (CL")

-

CDV Vertical s t a b i l i z e r drag c o e f f i c i w L

21 12 C

(CW)

Vertica s t a b i l i z e r angle of a t t a c k

21 13 C ALFVl rad

(- t a d (W/W))( o ( vl)

-

C cos (ALFv1) 21 14 CALFV1

_-

C SALFVl sin (ALFv1) 21 15 l b / f t2 Effective dynamic pressure a t v e r t i c a l 21 16 C QOV

s t a b i l i z e r (5")

DRP rad Rudder deflection corrected f o r

21 17 c

l a r g e deflection effects ( 6 r)

u Large deflection correction f a c t o r

21 18 C EKRIfD

f o r rudder (KEFF )

r

--

C Vertical s t a b i l i z e r p i t c h moment

21 19 cnv

c o e f f i c i e n t due to f l a p s (

w

A-34 Coded.

Segmnt Element Type Name Unite ikf i n i t i o n

21 20 C DXVT u Normalized Bo CV)8hift of vertical

rtabiliter cenfar of preasure as a

fuacticm of angle of a t t a c k ( A%)

21 21 C V T E F F - V e r t i c a l tail effectiveueea f a c t o r

C V T d

A.23 Segment 22 - haelage Aero Variables

Coded Semen t Elemeat Type Name Unite Definition 22 1 C I A X P l b X body axle fuselage aerodynamic

force ( A x , )

22 2 C D Y F l b Y body axis fuselage aerodynamic

force ( b y , )

22 3 C DZF l b 2 body axis fuselage aerodynamic

force (A+)

22 4

C DLF f t lb Fuselage body axis r o l l moment (ALF)

22 5 C D M F f t l b Fuselage body axis p i t c h moment ‘.A%)

22 6 C MIF f t l b h e e l a g e body axis yaw moment ( A N F )

22 7 c ALFF rad A i r c r a f t angle of attkck (a(), also

coded as - U H in segment 2Y and ALFG

In segment 20

22 a

C ALFYF rad Fuselage yaw angle of a t t a c k ( C J F ) 22 9 C A L F T F rad 22 10 C QXY l b / f t 2 Dynamic pressure based on UAS a d

VAS (= 1/2 p (UAS2 + VAS2)) Nq)

22 11

C QO l b / f t 2 Total dynlmfc pressure (7) , sleo

coded as QBAR in segment 29

A-35

Coded

Segrent Element Type Name &its Def i n i t f o n

22 12 C lblf t2

22 13 C f t2

22 14 C lb PLselage axial force (AF)

c

22 15 C sin (AF) - AF is el-t 17 in

segment 22

22 16 C f t 2 Fuselage aide force c o e f f i c i e n t (S/q)

rad Angle redefined tires while

22 17 C c a l c u l a t h g fuselage aero forces and m t e f t 2 Fuselage normal force coefficient 22 18 C

Wq)

hselage p i t c h m m t c o e f f i c i e n t

22 19 C f t

(nld

C f t Fuselage p i ch -rent c o e f f i c i e n t 22 20

Men AF =lf 5 /(18(ALP - AOF))

C l b Fuselage side force (Sp)

22 21

l b Fuselage normal force (Np) 22 22 C f t l b Fuselage p i t c h nok'ent about reference 22 23 C

cg (M)

-

C Not used 22 24 Fuselage yaw mment coefficient (N/q) C f t 3 22 25 f t Fuselage yaw -Pent vhen

22 26 C

AF = W2/(18(A2i. J )

Fuselage yaw norent about reference

22 27 C f t Ib

cg (N)

f t S h i f t of fuselage center of pressure 22 28 C as a function of angle of a t t a c k

( A m p )

A-36

A.24 mt 23 - D i r e c t Thrurt Force and lbrat V u l a b l a r

NUl'B: For a11 atrbmcriptd tmrlablu in sagcot 23, 1-1 Micatea left eaginc

and 112 ibdicatea right mint

b d e d

Segment Element 1Lw N & units Def iaitfoa

23 1 X body axla d i r e c t t h 8 t

C l b

force (Xml,

23 2 C lb P body mi8 d i r e c t t h r u s t

force ( ' - 9

23 C

3 l b 2 body axis d i r e c t t h r u s t

force (ZRs3

23 4 C f t / l b D i r e c t t h s t body axis r o l l

u a l t ( L d C 23 5 f t / l b D i r e c t tbrust body axis p i t c h -mt CHST)

6 C f t / l b D l r e c t thrust body axis yaw

. 23

-=t (IOTBST)

23 7-8 C rad P l a t turnlq angle of exhsust

(A#))

23 9-10 C l b N e t t h r u s t applied t o a i r c r a f t

a f t e r correction f o r flow

tirrnw ~ O S S U ~ (PCOB(1))

23 11-12 C l b X body axis force applied by I t h

engine (AX(I))

l b Y body axis force applied by I t h

23 13-14 C

engine ( A NI))

23 15-16 C l b Z body axis forceapplied by I t h

engine ( Aaz(1))

23 17-18 C f t X body axis distance between

nozzle swivel point and exit

plane ( A?S&))

P body axis distance between

23 19-20 C f t

nozzle swivel point and e x i t

plane ( A BLw(I))

23 21-22 C f t 2 body axis distance between

nozzle swivel point and exit

plane ( A nsw(1))

-

23 23-24 C sin (PSIT(1))-PSIT(1) located

in segment 3, elements 24 and 25

-

cos (PSIT(1))-PSIT(1) located

23 25-26 c

in segment 3, elements 24 and 25

-

27-28 C sin (Ttrr(1)) THT(1) located i n

segment 3, elements 26 and 27

-

23 29-30 C cos (TRT(I))-l"T(I) located in

segment 3, elements 26 and 27

-

Flow p r n l n g correction factor 23 31 C A-37

A.2S 24 - Inlet Ram Force and llament Variables

BOTE: ?or all mbocriptd v u l a b l u in - t 24, 1 1 1 iadicattu left emgiPI, 1-2 indicatu raht -be.

l b X body axis inlet ram force

24 1 C

'h'

1 body axis inlet r a m force

2 C lb

Z bod7 u l a i n l e t ram force

l b

24 3 C

(%An) ft l b Inlet ram body axis r o l l moment

4 C

(b)

Inlet ram body axis pitch moment

f t I b

24 5 C

%'

it lb Inlet ram body axis yaw moment

C

24 6

%An'

True airs@ (PA), also cod&

7 C f t l s c c as VA in segment 29 Geometric inlet flaw turning 8 C rad

angle in X-2 I n l e t plane (hilar )

Geometric inlet flow turning angle

c rad

in ii-Y inlet plane (B-)

-

cos (BTURN)

C 24 10

-

sin (BTURN)

1 1 C I n l e t f l o w v e l o c i t y (Vml

12 C f t/8=

-

Ratio of true airspeed t o inlet 13-14 C flow velocity ( V O v & ) ) Actual i n l e t flow turn- angle rad 15-16 C

in X-2 inlet plane (ATv&))

-

C I n l e t ran monent arm normalized 24 17-18 t o eauivalent i n l e t diameter MTURN(1) rad Difference between a c t u t l and 24 19-20 C geometric I n l e t flow turn-

angles ( A Am&))

ETAR(I) - Ratio of actual t o t h e o r e t i c a l inlet

C 24 21-22

ram forre q i t u d e (q R(I)